Development and Experimental Verification of a Hybrid VerticalTake-Off and Landing (VTOL) Unmanned Aerial Vehicle(UAV)

Haowei Gu, Ximin Lyu, Zexiang Li, Shaojie Shen, and Fu Zhang

Abstract— In this paper, we present the design and ver-ification of a hybrid vertical takeoff and landing (VTOL)unmanned aerial vehicles (UAV) of the type named dual-system or extra propulsion VTOL UAV. This paper features thecomprehensive system development of such VTOL UAVs fromall aspects, including the aircraft design and implementation,onboard devices integration, ground station support, and longdistance communication. We proceed with aerodynamic anal-ysis, mechanical design, and controller development. Finally,we verify by experiment that this hybrid VTOL UAV has thedesired aerodynamic performance, flight stability, enduranceand range. In addition, with the designed flight controller, theVTOL UAV can achieve full autonomous flight in a real outdoorenvironment. It serves a good platform for future research, suchas vision-based precise landing, motion planing and quick 3-D mapping, as well as service applications, such as medicinedelivery.

I. INTRODUCTION

During the last decades, unmanned aerial vehicles (UAV)have been widely used in both military and consumer areassuch as surveillance, tracking, monitoring and aerial pho-tograpy. These UAVs are mainly classified into two types:fixed-wing UAVs and rotary-wing UAVs and each type hastheir own advantages and drawbacks. As for fixed-wingUAVs, they have several features, such as high cruising speedand altitude, high flight efficiency and large cabin, whichmeans they can fulfill many tasks that require the vehicle tocarry a heavy payload or achieve long endurance and range.However, fixed-wing UAVs have their limits. A fixed-wingairplane needs a runway or catapult to take off and land,which imposes restrictions on its applications. Rotary-wingUAVs, on the other hand, can take off and land vertically,which makes these vehicles suitable for a wide variety ofenvironments. But rotary-wing vehicles do not have longendurance or range. It is still a challenge to develop a small-scale UAV, which is able to achieve long flight enduranceand range and at the same time takes off and land vertically.Hybrid vertical takeoff and landing (VTOL) UAVs are asuitable solution for this problem; they combine the conceptsof fixed-wing and rotary-wing UAVs in a single platform,thus inheriting the advantages of both.

There are several types of hybrid VTOL UAVs, includingtail-sitter, tilt-rotors and dual-system UAVs.

*Research supported by Hong Kong ITF Foundation (ITS/334/15FP).All authors are with the Department of Electronic and

Computer Engineering, Hong Kong University of Science andTechnology, Hong Kong, China. hguad@connect.ust.hk,xlva@connect.ust.hk, eezxli@ust.hk,eeshaojie@ust.hk, eefzhang@ust.hk

Fig. 1. The developed hybrid VTOL UAV (Coordinate systems aredefined in section III) Video available at: https://www.youtube.com/watch?v=wRE4ICjHJ_0

A. Tail-sitter VTOL UAVs

A tail-sitter VTOL UAV takes off vertically on its tail, thenthe whole aircraft including the fuselage tilts horizontallyfor level flight. In [1] and [2], many aspects of a tail-sitter have been studied, including design, manufacturing,modeling, control and flight tests. However, since the samemotors are being used for both hovering and level flight, afixed-pitch propeller is usually not able to achieve optimalefficiency in both stages while a varying-pitch propeller cansignificantly complicate the mechanical design, maintenanceand manufacturing cost. In addition, the big wing area beingexposed to cross-wind during hovering introduces a con-siderable disturbance that is likely to degrade the hoveringperformance.

B. Tilt-rotor VTOL UAVs

Similar to a tail-sitter VTOL UAV, a tilt-rotor VTOL UAVgenerates lift and forward thrust using the same propulsionsystem. But instead of tilting the whole aircraft like a tail-sitter VTOL UAV, the tilt-rotor VTOL UAV only tilts itsrotors. For vertical flight, such as takeoff, hovering andlanding, rotors of the propulsion system are angled suchthat their plane of rotors is horizontal, lifting in the waythat a helicopter rotor does [3]. As the UAV gains speed,the rotors are progressively tilted forward, with the planeof the rotors eventually becoming vertical. In this mode,the wing provides the lift, and the rotor provides thrust

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as a propeller. In [4], [5] and [6], the structure design,aerodynamic analysis and flight control were addressed. [7]shows another configuration of tilt-rotor, which also calledtilt-wing VTOL UAV. The main drawback of a tilt-rotorVTOL UAV is its complicated structure, which increases theaircraft weight and cost.

C. Dual-system VTOL UAVs

A dual-system VTOL UAV has two independent propul-sion systems for hover and level flight. As seen in Fig. 2, theaircraft takes off vertically using the lift force generated bythe four lifting rotors. During this phase, the push motorwill be turned off and the whole aircraft behaves like aquadrotor [8], its position as well as attitude are controlledby the differential thrust of the four rotors. After reachingthe desired altitude, the aircraft starts transiting to level flightby turning on the push motor to gain forward speed. Oncethe cruise speed is reached, the four lifting rotors will beturned off and the whole aircraft behaves like a conventionalfixed-wing airplane. The control of the position and attitudeis achieved by the push motor and control surfaces suchas the ailerons, elevators and rudders. Since the propulsionsystems for hover and level flight are independent, they canbe optimized for both flight modes separately. Other thanmotors and propellers, all parts of a dual-system VTOL UAVare assembled rigidly and no moving parts are involved. Asa result, they are easy to design, manufacture and maintain.

Aileron

Lifting Motor

Vertical Tail

Horizontal TailElevator

Wing

Pusher Motor

Fig. 2. Hybrid VTOL UAV explanation

In this paper, we present the development of a dual-systemhybrid VTOL UAV from a holistic system design perspective.The primary goal of the designed aircraft is to deliver alightweight parcel, such as first aid medicine, at a distance ofover 30 km, while maintaining a relative small size for betterportability. The designed unmanned aerial system has all thekey components for safe and reliable operation, includingautonomous navigation, robust flight controller, first-person-view (FPV) camera, long distance communication, groundstation, etc. The detailed aerodynamic design, analysis, me-chanical consideration and flight controller development are

presented. The designed aircraft is shown in Fig. 1. Asvalidated by real flight tests, the developed UAV can achievethe required flight range and fully autonomous operation.

The remainder of this paper is organized as follows. Theaerodynamic and mechanical design, as well as onboarddevices integration, are detailed in section II. Position con-trollers, attitude controllers and VTOL mixer are presented insection III. Section IV presents the details of the experimentsetup and results discussion. And finally section V draws theconclusion.

II. SYSTEM DESIGN

A. Airframe and Propulsion System Selection

As mentioned before, the designed aircraft is expected toachieve a flight range of 30 km. For this purpose, a varietyof devices need to be used, including video transmitter foraircraft status monitoring, telemetry for long distance datacommunication, radio control (RC) extender for emergencypiloting, etc. Their estimated weights are summarized intable I, where the units are in grams.

TABLE ITABLE OF DEVICE WEIGHT

Device Quantity Weight1

Video transmitter 1 20Telemetry 1 20

Battery 1 900Flight control board 1 30

Lifting motors 4 70ESCs for lifting motor 4 40

Pusher motor 1 80ESC for push motor 1 60

Others N.A. 200Total weight N.A. 1750

The item ”Others” includes servo motors, wires, pitot-tube, GPS, etc. The total weight estimated for all of thedevices is around 1.75 Kg . In our work, we intend toadopt a commercially available fixed-wing airplane as thestarting point to develop a hybrid VTOL UAV due to theirproved flight stability and performance. Table II shows thekey configurations of a number of commercial-off-the-shellfixed-wing airplanes.

TABLE IIFIXED-WING AIRPLANE MODELS

UAV MiniSkyHunter

Sky-Hunter Talon Ranger

EX757-4Wing Span (mm) 1238 1800 1718 1380

Gross Weight (gram) 760 900 1050 750MTOW (gram) 2500 3500 3000 1100

Payload Capacity(gram) 1740 2100 1950 350

It can be seen that Mini Skyhunter has the smallest sizewhile achieving the required payload capacity. A smaller size

1The unit of weight is in gram

161

is preferable so to increase its maneuverability and ease ofthe flight test. The detailed parameters of such an airplaneare summarized in table III, where mean chord, wing areaand aspect ratio (AR) are measured on a real airplane. Thecruise speed is recommended by the manufacturer. e is theOswald efficiency factor which is typically between 0.7 and0.85 and can be computed from AR [9].

TABLE IIIDETAILED PARAMETERS OF MINI SKYHUNTER

Mean AerodynamicChord (MAC) Wing area Cruise speed AR e

17 cm 0.21 m2 15 m/s 7.28 0.83ρ MTOW Re CL CD

1.157 kg/m3 2500 gram 2.3× 106 0.896 0.064

To compute the Reynolds number Re, the commercialsoftware Profili is used. Supplying the airspeed; characteristiclinear dimension, which is defined as the mean aerodynamicchord (MAC); and cruise altitude, which is 100 m for ourapplication to the software, yields the Reynolds numberRe = 2.3× 106.

Lift coefficient CL [9] is calculated for level flight wherethe maximal takeoff weight MTOW = 24.5 N and cruisespeed V = 15 m/s. As pointed out in [10], the total drag ofa typical airplane consists of three parts: parasitic drag D0,wave drag Dw and induced drag Di. The wave drag occurswhen the airplane’s speed is near the speed of sound and canbe safely ignored in our case where the flight speed is farfrom the speed of sound. As shown in [11], the total drag,when parameterized by drag coefficient, can be expressed as

CD =Swet

S· Cfe +

C2L

π ·AR · e(1)

where the typical value of Swet/S for the Mini SkyHunteris around 4, as seen in [12], and the skin friction coefficientCfe is drawn from [13].

After determining the aerodynamic configuration of theaircraft, we proceed to the propulsion system design. Fora hybrid VTOL UAV, there are two sets of propulsionsystems, namely the four lifting rotors for takeoff, landingand hover flight, and the push motor for level flight. TableIV summarizes the key parameters of a few motor-propellerpairs specifically designed for multirotors.

TABLE IVPARAMETERS OF MOTORS AND PROPELLERS FOR MULTIROTORS

Name KV Thrust (gram) Propeller Voltage (V)DJI E310 960 800×4 9×4.5 11.1DJI E800 350 2100×4 13×4.5 22.2

EMax MT2212 900 850×4 9×4.5 14.8EMax MT2208 1500 690×4 6×45 14.8

It is clear that the thrust produced by the four liftingrotors has to be greater than the force of gravity acting onthe vehicle. To keep a sufficient margin for maneuvering,

the thrust/gravity ratio is designed to be 1.3. The resultingmaximal thrust of each motor is therefore 0.81 kg, makingDJIE310 a good choice.

In a similar way, thrust produced by the push motor isdesigned to be 2 times greater than the drag force in order togain sufficient maneuverability, such as during acceleration,banking turns, and level flight. That is:

T = 2 ·D = 2× 1

2ρV 2SCD

= 2× 0.5× 1.157× 152 × 0.21× 0.064 = 3.498N.(2)

With this requirement, we investigate a few motors andpropellers designed for fixed-wing airplanes, as summarizedin tableV. It can be seen that the Sunny Sky X2216 is theright fit for to our needs.

TABLE VPARAMETERS OF MOTORS AND PROPELLERS FOR FIXED-WING

AIRPLANE MODELS

Name KV Thrust(N ) Propeller Efficiency

(g/W )Voltage

(V )SunnySky X2826 880 4.9 12×6 0.068 14.8SunnySky X2216 880 3.92 10×4.7 0.072 14.8SunnySky X2814 900 4.9 11×7 0.070 14.8

B. Mechanical Design

Even though the main airframe is adopted from acommercial-off-the-shell airplane, the integration of fourlifting rotors for vertical takeoff and landing introduce newproblems that need to be carefully considered. A majorchallenge is the structural design and precise landing in

Fig. 3. motor layout and inclination

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the presence of cross wind. The four lifting motors arelocated around the fuselage systematically along the X axisand Y axis, as shown in Fig. 3. The vertical tails of themain airframe, primarily designed for enhancing the direc-tional stability of the aircraft during level flight, introducea considerable aerodynamic moment in the yawing duringhover flight in the presence of cross wind. This aerodynamicmoment, together with other aerodynamic disturbances fromthe wing and fuselage, will affect the hover accuracy andlanding reliability. In order to mitigate this situation, a largeyawing moment that usually exceeds the reaction torque ofa typical motor-propeller pair is required. As such, the fourlifting motors of the vehicle are inclined along their diagonallines by a small angle δ. As shown in Fig. 3, the green circlearrow indicates the direction of the motor reaction torque,and the red arrow represents the inclination direction. Withthis design, the total moment in yawing is the sum of thetorque produced by the motor thrust projected to the bodyX − Y plane and the motor reaction torque. In addition, asthe motor inclination is along the motor diagonal lines, thetorque can be maximized with the same motor thrust andinclination angle.

C. Others

To constitute a comprehensive unmanned aerial systemthat is able to perform a mission fully autonomously andwith the capability of communicating with a ground stationin real time to report its flight status, such as air data, positionand battery status, as well as for receiving commands fromthe ground station in case of an emergency or mission re-planning, several onboard devices need to be integrated.After investigating the current commercial-off-the-shelf de-vices, we have settled on the devices given in table VI.

As the designed aircraft is expected to achieve a flightdistance of 30 km, to support its long distance operation, anautomatic antenna tracker (AAT) is used extend the connec-tion between the aircraft and the ground station. An AAT cantrack the aircraft using its GPS signal, by way of an antennainstalled on the ground station that can always point to theaircraft. By doing so, the data link can be greatly enhanced.Telemetry is used to connect the vehicle and ground stationto download flight status and upload commands sent fromthe ground station in real time. A lightweight first personview (FPV) camera and the associated video transmitter areused to monitor the aircraft’s surrounding environment inreal time. A radio extender is used to extend the operationrange of the radio controller. The flight controller used isa Pixhawk autopilot open hardware and software platform[14]. The Pixhawk autopilot has integrated the necessarysensors for autonomous navigation, such as a magnetometerfor measuring the direction of south, an accelerometer formeasuring the direction of gravity, gyros for measuringangular rate and a barometric altimeter for measuring thealtitude of the aircraft. With this open source platform, wedesigned our own flight controller, as discussed in nextsection, in detail. Other key components such as a pitot-tube, GPS and data transmitter are also used although are

TABLE VISELECTION OF DEVICES

Name Figure Brand Frequency

Communi-

cation

range

AATMy Fly

DreamN.A.

Depands

on

vedio

transmitter

Telemetry XTend 915MHz 32 Km

Video

transmitterLawMate 1.4GHz 35Km

Radio

extenderU Link 433M 40 Km

Camera ANAITV N.A. N.A.

Pixhawk 3DR N.A. N.A.

not shown in table VI.Fig. 4 shows the entire unmanned aerial system that

has been implemented. The whole system consists of aground station, a pilot, and the unmanned aerial vehicle. Theground station is a PC running Qgroundcontrol (QGC) [15]software for logging and displaying the aircraft status, anAAT for aircraft tracking and connection enhancement, andan external monitor that displays the onboard FPV video aswell as the flight status of the aircraft such as the airspeed,distance, altitude, speed and attitude. A pilot can maneuverthe aircraft directly via a radio controller if in manual testingmode, or simply monitor the aircraft status and take over itscontrol in case of emergency if in fully autonomous flightmode.

III. FLIGHT CONTROL SYSTEM DESIGN

The block diagram of our flight control system is shown inFig. 5. It can be seen that the whole flight system consists ofthree level controllers: position controller, attitude controller,and VTOL mixer.

Comprising the first level are three position controllers:a transition controller, rotary-wing position controller, andfixed-wing position controller, which are running simultane-ously. A position controller receives the desired position from

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FPV Video

GPS Signal

Aircraft status

Manual Command

Enhanced Manual

Command

QGC Command

Radio Controller Ground Station Monitor AAT

Vehicle with Autopilot

Radio Extender

Fig. 4. Signal flow of whole unmanned aerial system

either human pilots or navigation way points and computesthe desired attitude and thrust, which are then fed to asignal distribution module. The signal distribution modulemaps the desired attitude from the transition controller aswell as, rotary-wing position controller to the same attitudecontroller, called the VTOL attitude controller, and mapsthe attitude from fixed-wing position controller to the cor-responding fixed-wing attitude controller. The thrust of thethree position controllers are fed to a VTOL mixer for furtherprocessing.

The second level controllers are two attitude controllers: afixed-wing attitude controller and VTOL attitude controller.While the fixed-wing attitude controller directly computesthe desired moment from its attitude error, the VTOL atti-tude controller will use either the transition or rotary-wingposition controller attitude set point to compute the desiredmoment, depending on the current flight mode.

Depending on the current vehicle flight mode (i.e., rotary-wing mode, fixed-wing mode or transition mode), the VTOLmixer chooses different moment and thrust sources and mapsthem to different actuators. For example, in rotary-wingmode, the moment from the VTOL attitude controller andthe thrust from the rotary-wing position controller are usedto actuate the four lifting rotors. This results in an aircraft thatwill behave like a typical quadrotor UAV. In transition flightmode, the moment from the VTOL attitude controller and thethrust from the transition controller are synthesized to actuatethe control surfaces such as ailerons, push motor, or the fourlifting rotors all together. More specifically, during forwardtransition from rotary-wing mode to fixed-wing mode, thepush motor speeds up to a prescribed RPM to provideforward thrust. Once the prescribed airspeed is reached,the fixed-wing mode will be triggered. During backwardtransition from fixed-wing mode to rotary-wing mode, thepush motor slows down gradually to a stop. At the sametime, the lifting rotors are turned on to control the vehicle’sattitude and altitude. Similarly, once the vehicle’s airspeeddrops below some prescribed value, the rotary-wing modewill be triggered. Finally, in fixed-wing mode, moment from

fixed-wing attitude controller and thrust from the fixed-wingposition controller are used to actuate the control surfacesand push motor like a conventional fixed-wing airplane.

A. Position and Attitude Controller Implementation

In our scenario, the position controller including fixed-wing position controller, rotary-wing position controller, andtransition controller. The fixed-wing position controller isadopted from [16] and the rotary-wing position controllerfrom [17]. The transition controller will be presented insection III-B. The fixed-wing attitude controller is a PIDcontroller adopted from [18]. The VTOL attitude controlleris a cascaded controller similar to [1]. These controllers areall implemented on the existing framework of the Pixhawkautopilot software stack [18].

B. Transition Controller

Here we focus on introducing the transition controller. Thekey features of the transition controller in Fig. 5 are sendingfixed attitude command, the pitch and roll commands areboth set to 0, while the yaw is set to its current value at thebeginning of transition. The push motor is set to its maximumpower until the specified airspeed is reached.

The thrust command is determined by an altitude holdingcontroller, which is used to hold the aircraft altitude at thevalue right before the transition. Recall that the rigid bodyaltitude dynamic can be written as

hz =1

mrT3 (Faero + FT ) + g. (3)

where r3 denotes the third column of the rotation matrixR ∈ SO(3), which is the orientation of the body framerelative to the inertial frame (seen in Fig. 1), hz is thealtitude in the inertial frame, Faero ∈ R3 and FT =[FTx FTy FTz

]T ∈ R3 are the aerodynamic force andmotor thrust vector in body frame, respectively.

The reference altitude is described as hd, and the altitudeerror is described as he = hd − hz . A simple altitude PIDcontroller is used to calculate motor thrust as follows:

rT3 FT +mg = kphe + ki

∫hedt+ kd

dhe

dt. (4)

The resulting thrust can be finally described as

FTz =kphe + ki

∫hedt+ kd

dhe

dt −mg

r31. (5)

where r31 is the first entry of r3 ∈ R3. It is worth noticingthat the aerodynamic force Faero, though neglected in thecontroller design, can be compensated effectively by theintegral action. The computed thrust and attitude commandsare sent to both VTOL and fixed-wing attitude controller formoment computation. A VTOL Mixer, as discussed below,is finally used to distribute the required moment and thrustto proper actuators (i.e. control surfaces and lifting motors)depending on the current airspeed during transition.

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Fixed-wing

Position Controller

Navigation

Way Points

Or

Manual Inputs

Rotary-wing

Position Controller

VTOL Attitude

Controller

VTOL

Mixer Aircraft

Dynamics

Signal

Distribution

Reference position

Moment_1

PWM

Position error

Vehicle Position

Vehicle Attitude

+ _

Transition

Controller

Thrust_1 Att_sp1

Fixed-wing

Attitude Controller

Thrust_1&Thrust_2&Thrust _3

Thrust_2 Att_sp2

Thrust_3Att_sp3

Att_sp1Att_sp2

Att_sp3

Flight Mode

Moment_2

Fig. 5. Architecture of the flight control system

l1δ1 δ2

δiδ3

f2f1yb

xbb

Fig. 6. Vehicle configuration and effect of elevons

C. VTOL Mixer

The available actuators of our vehicle are shown in Fig. 6.δ1 and δ2 are the angle of the left aileron and right aileron,respectively. δ3 is the angle of the elevator. f1 and f2are the lift forces produced by the left wing and rightwing, respectively. l1 is the distance between f1, which actson the mean aerodynamic chord (MAC) and longitudinalaxis. Because of symmetry, the distance between f2 andlongitudinal axis is l1 as well. According to [19], f1 andf2 can be written as

f1 =1

2ρV 2Sl (CL0 + CLαα+ CLδ

δ1) .

f2 =1

2ρV 2Sr (CL0 + CLαα+ CLδ

δ2) .(6)

where ρ is the air density, V is the airspeed, and Sl andSr are the left and right wing area, respectively. Because ofsymmetry, Sl and Sr are the same. α is the attack angle.CL0

is the lift coefficient while α = 0, and δi = 0. CLα

and CLδare, respectively, the lift coefficient derivatives with

respect to attack angle and aileron deflection. Similarly, theaerodynamic pitch moment can be written as

Maero =1

2ρV 2S (Cm0 + Cmαα+ Cmδ

δ3) c. (7)

where S is the total area including tails, c is the mean aero-dynamic chord (MAC), Cm0

is the pitch moment coefficientwhen α = 0 and δi = 0. Cmα and Cmδ

are, respectively, thepitch moment derivatives with respect to attack angle andelevator deflection. As no rudder is used for this aircraft,the moment produced by all control surfaces can be furtherwritten as [

2MTx

ρV 2Sll1CLδ2MTy

ρV 2ScCmδ

]=

[1 −1 00 0 −1

]δ1δ2δ3

δ1 = δ2

(8)

where α1 and α2 are two different first order functions thatrelate to airspeed. The moment and thrust produced by thefour lifting rotors can be represented as:

FTz

cos(δ)MTx

sin(δ)d+κ cos(δ)MTy

(cos(δ)d−κ sin(δ)) sin(χ)MTz

(cos(δ)d−κ sin(δ)) cos(χ)

=

1 1 1 1−1 1 −1 11 1 −1 −1−1 1 1 −1

T1

T2

T3

T4

(9)

and

Ti = ct$2i (i = 1, 2, 3, 4) . (10)

κ =cqct. (11)

where d is the length of each motor to the aircraft center ofmass, as seen in Fig. 6 and χ is the azimuthal angle of thefirst rotor with respect to the lateral axis of the aircraft. δ isthe inclination angle of the four lifting rotors, as presentedin the previous section. Ti is the rotor thrust and $i is therotation speed of each rotor (i = 1, 2, 3, 4). ct and cq are thepropeller thrust coefficient and propeller torque coefficient,respectively. κ is the ratio of torque to thrust coefficient.

It can be seen that the roll and pitch moment can beproduced by either control surfaces, as shown in Eq. (8)or by motor differential thrust, as shown in Eq. (9). In ourimplementation, we assign a weighting factor α to weightthe contribution of the moment from the control surface andthe motor differential thrust. That is, given the required rollmoment L, the moment from the control surface and themotor differential thrust are, respectively, (1−α)L and αL.The way to allocate pitch moment is similar. The weightingfactor α is in between zero and one, depending on the

165

current flight mode. In fixed-wing mode, α = 0, and therequired moments are solely generated by control surfaces,the actuation signal of which are computed by inverting Eq.(8). Accordingly, the required thrust is mapped to the pushmotor. In rotary-wing mode, α = 1, the required moment aswell as thrust are solely generated by the four lifting rotors,the actuation signals of which are computed by inverting Eq.(9). In transition flight, the required moments are generatedby the control surface with portion 1 − α and by motordifferential thrust with portion α. α = 1− (0.6 V

VT+ 0.2) is

a function of airspeed V , indicating that the effectiveness ofthe control surfaces increases as the airspeed V increases.VT is a barrier speed that the aircraft has to exceed inorder to switch to full fixed-wing mode. The total thrustproduced by the four lifting rotors is computed from thealtitude holding controller presented previously, while thepush motor is set to a prescribed PWM to accelerate theaircraft in a horizontal direction. In practice, the parametersin Eq. (8) and Eq. (9) are not known, resulting a gainmismatch when inverting these two equations. By properlytunning the parameters of the flight controllers, these gainmismatches can be effectively mitigated, as shown in realflight tests.

IV. EXPERIMENTS

With the designed dual-system VTOL UAV along with itsflight control system, intensive experiments are conductedto verify its aerodynamic performance, flight stability andsystem level autonomy. The first experiment is designedto verify the reliability of transition controller, the secondexperiment is to test the flight controller’s stability when theaircraft is under active pilot control, the third experiment isto test its range, endurance and the capacity of autonomousflight, and the fourth experiment is to determine the optimalflight speed.

A. Transition controller verification

Since the transition is a necessary maneuver of the hy-brid VTOL UAV to achieve both vertical takeoff and levelflight, we designed an experiment to verify the reliability ofthe transition controller presented in previous sections. Theprocedure of this experiment is as follows: the aircraft firsttakes off in manual mode. After reaching the desired altitude,the aircraft was switched to autonomous flight mode, wherethe quadrotor position controller took place and controlledthe aircraft to a transition point. Once the transition pointis reached, the aircraft initiated a transition maneuver andsubsequently flied as a fixed-wing airplane.

Fig. 7 and 8 show the transition process. The upper figurein Fig. 7 indicates the flight states where 0 means quadrotormode, 1 means transition mode and 2 means fixed wing flightmode. As mentioned in previous sections, after the transitionmaneuver is initiated, the push motor PWM increases linearlyuntil the maximal value is reached, resulting an increasedairspeed. During the transition process, the altitude increasedby 1m due to the increased aerodynamic lift. After theairspeed reached the specified value, the flight state switched

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.5828

30

32

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.58

1000

1500

2000

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.58

0

1

2

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.580

20

40

Fig. 7. Flight status, throttle, airspeed and altitude during transition process

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.58-50

0

50

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.58-20

-10

0

10

1.42 1.44 1.46 1.48 1.5 1.52 1.54 1.56 1.58-50

0

50

100

Actual pitch anglePitch angle set point

Actual roll angleRoll angle set point

Actual yaw angleYaw angle set point

Fig. 8. Roll, pitch and yaw angle during transition

to fixed wing mode, and the push motor is controlled bya typical fixed-wing flight controller. Fig. 8 shows attituderesponse during the transition. It is worth noticing that theroll and yaw angles are set to relative large values after thetransition due to the bank turn maneuvering. This resultsa slight altitude loss (i.e. 2m) and attitude error, whichare caused by the low bandwidth of a typical fixed-wingcontroller. These errors, in practice, are acceptable in ouroutdoor environment.

166

B. Manual Flight Test

As the whole vehicle is a new system whose aerodynamicperformance, robusteness and system reliability remain tobe verified, we designed a manual flight experiment for thepurpose of determining the minimal airspeed, verifying theposition controller and each chosen component of the wholesystems such as video transmitter, telemetry, and AAT.

This experiment is designed as follows: First, the hybridVTOL UAV takes off vertically in manual mode and lifts upto about 50 m. Then, the test pilot triggers a switch commandto level flight mode. After the vehicle has fully transited tothe fixed-wing mode, the test pilot will control the aicraftsuch that it can approximately track a circle. Finally, thehybrid VTOL UAV is switched to quadrotor mode and islanded.

0.8 1 1.2 1.4 1.6 1.8 2 2.214

14.5

15

15.5

16

16.5

AirspeedFitting curve for airspeed

0.8 1 1.2 1.4 1.6 1.8 2 2.244

45

46

47

48

49

Altitude masured by GPSSet point of altitude

Fig. 9. Altitude and airspeed of manual flight test

0.8 1 1.2 1.4 1.6 1.8 2 2.2-50

0

50

0.8 1 1.2 1.4 1.6 1.8 2 2.2-10

0

10

0.8 1 1.2 1.4 1.6 1.8 2 2.2-200

0

200

Fig. 10. Roll, pitch and yaw angle of manual flight test

0

160

14080

12060

10040

802060

040

-2020

-400

-20 -60

20

40

Fig. 11. Trajectory of manual flight test

The data log for circling flight is shown in Fig. 9, Fig. 10and Fig. 11. The upper figure in Fig. 9 shows the altituderesponse. It can be seen that the actual altitude is wellregulated around the desired value, with about 2.6 m peakto peak oscillation. This oscillation, in the same period ofthe flight trajectory, is caused by cross wind and lead to asimilar oscillation in airspeed, as shown in the lower plot inFig. 9. Fig. 10 shows the roll, pitch and yaw angle of theaircraft in manual mode where the attitude angle is directlycontrolled by pilot. Due to the presence of cross wind, theaircraft slowly drifts from its initial trajectory, the pilot thusintentionally adjusted roll angles respectively at 1min, 1.2minand 2min to maintain its trajectory within the proper area.The resulting trajectory is shown in Fig. 11.

C. Fully Autonomous Flight and Flight Range Verification

As pointed out in section I, our primary goal is todevelop a UAV which can take-off and land vertically andachieve a 30 km range in a fully autonomous way. Thus,this experiment is designed to verify whether our UAV canachieve this goal.

Due to the limited flight test area, we cannot let the vehiclefly straight forward for 30-kilometers. The compromisedsolution is to follow a circle trajectory and calculate thedistance by airspeed and endurance time. In this experiment,we set the airspeed as 18 m/s (64.8 km/h). Then a minimalflight endurance of 27.78 min is required to meet 30 km ofrange. To verify this, a fully autonomous flight is performedon our unmanned aerial system, including auto take-off,transition and level flight.

The log data is shown in Fig. 12, Fig. 14, Fig. 13 and Fig.15. In Fig. 12, the upper plot shows the aircraft airspeed,which oscillates in the same period of flight trajectory dueto the presence of cross wind. This results an oscillationin current measurement (middle plot of Fig. 12), roll angle(upper plot of Fig. 13) and pitch angle (middle plot ofFig. 13). Moreover, it is seen that our vehicle can fly32 min, satisfying the required endurance (i.e. 27.78min).Fig. 14 shows the positions. It is noticed that the altitude is

167

0 5 10 15 20 25 30 35

4

6

8

10

12

14

0 5 10 15 20 25 30 3513

14

15

16

17

0 5 10 15 20 25 30 35

16

18

20

22

Fig. 12. Airspeed, current and voltage range test

oscillating, which is, again, caused by the cross wind andin the same period of flight trajectory. Fig. 15 shows theresulting trajectory.

0 5 10 15 20 25 30 3520

30

40

50

0 5 10 15 20 25 30 35-10

0

10

0 5 10 15 20 25 30 35-200

0

200

Fig. 13. Roll, pitch and yaw angle of range test

D. Optimal Flight Speed

Although the autonomous flight test presented above hasalready realized our goal, the range and endurance of the

0 5 10 15 20 25 30 35-100

0

100

200

0 5 10 15 20 25 30 350

100

200

0 5 10 15 20 25 30 35

25

30

35

Fig. 14. Position of range test

Fig. 15. Trajectory of range test

designed aircraft can be further improved by observing thefollowing:

• From Fig. 13 we can know that in the flight range test,the vehicle has a constant banking angle of about 33.7deg. This means that only 100% · cos 33.7◦ = 83.2%aerodynamic lift has been used to counterweigh theaircraft’s force of gravity. The flight range for straightlevel flight will be 1.2 times more.

• As we know, flight speed is cruicial in determining theflight efficiency. As for our VTOL, 18 m/s may not bethe optimal flight speed. Consequently, it is possible toextend the flight range if the aircraft flys at an optimalspeed.

To find the optimal flight speed, another experiment isconducted. Similar to the second experiment, the aircraft isset to loiter around a specified point at different test speeds.Figure 16 shows the current of each case which oscillateddue to disturbance. The average current and voltage aresummarized in table VII.

168

0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.22468

1012

16 m/s

0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.22468

1012

17 m/s

0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.22468

1012

18 m/s

0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.22468

1012

19 m/s

Fig. 16. This hybrid VTOL UAV is test in different airspeed to measurethe current to determine cruising speed

TABLE VIIAIRSPEED TEST RESULT

Airspeed V (m/s) 16 17 18 19Average Current I (A) 5.81 5.73 7.07 8.04

V/I 2.75 2.97 2.54 2.36

The expected range can be predicted by

R = t · V =C

I· V =

V

I· C. (12)

where R is the flight range, t is endurance time, C is batterycapacity and I is the current drained from the battery. Byassuming that the battery capacity C is constant, the flightrange is proportional to the ratio of airspeed V to currentI . As seen in table VII, the ratio V/I is maximal whenthe flight speed is set at 17 m/s, the corresponding flightrange will be 2.97

2.54 = 1.17 higher than the case where V =18m/s. Therefore, the flight range our aircraft can achieveis expected to be 1.2× 1.17× 30 km = 42 km.

V. CONCLUSION AND FUTURE WORK

In this study, a hybrid VTOL UAV was developed andtested. The design requirement was achieving a flight rangeof 30 km. The work is based on a mini SkyHunter due toits size and payload capacity. On the basis of theoreticalcalculation, selection of the propulsion system and mechan-ical design proceeded. A flight control system was designedto realize transition, improve flight stability and achieve

full autonomous flight. The designed flight control systemconsists of VTOL attitude controller, transition controllerand VTOL mixer based on our special aerodynamic char-acteristics and mechanical design. Finally, experiments weredesigned to determine the aerodynamic performance, verifythe controller’s stability and autonomy. As shown by theseexperiments, this prototype can meet the initial design goal.

In the future, we will optimize this prototype, in termsof aerodynamics design and controller improvement. Actuallong distance experiments will be conducted to verify therange of the communication devices. Based on this hybridVTOL UAV, several topics, such as precise landing, motionplanning, and 3-D mapping will be investigated.

REFERENCES

[1] X. Lyu, H. Gu, Y. Wang, Z. Li, S. Shen, and F. Zhang, “Designand implementation of a quadrotor tail-sitter vtol uav,” in 2017 IEEEInternational Conference on Robotics and Automation (ICRA). IEEE,2017(Accepted).

[2] F. Zhang, X. Lyu, Y. Wang, H. Gu, and Z. Li, “Modeling and flightcontrol simulation of a quadrotor tailsitter vtol uav,” in AIAA Modelingand Simulation Technologies Conference, 2017, p. 1561.

[3] R. W. Prouty, Helicopter performance, stability, and control, 1995.[4] R. T. Rysdyk and A. J. Calise, “Adaptive model inversion flight control

for tilt-rotor aircraft,” Journal of guidance, control, and dynamics,vol. 22, no. 3, pp. 402–407, 1999.

[5] G. Flores, I. Lugo, and R. Lozano, “6-dof hovering controller design ofthe quad tiltrotor aircraft: Simulations and experiments,” in Decisionand Control (CDC), 2014 IEEE 53rd Annual Conference on. IEEE,2014, pp. 6123–6128.

[6] D. Wu, H. Li, and S. Li, “Flight dynamics study of a small tilt rotoruav with tail propeller,” in 16th AIAA Aviation Technology, Integration,and Operations Conference, 2016, p. 3450.

[7] E. Cetinsoy, S. Dikyar, C. Hancer, K. Oner, E. Sirimoglu, M. Unel,and M. Aksit, “Design and construction of a novel quad tilt-wing uav,”Mechatronics, vol. 22, no. 6, pp. 723–745, 2012.

[8] S. Bouabdallah, P. Murrieri, and R. Siegwart, “Design and controlof an indoor micro quadrotor,” in Robotics and Automation, 2004.Proceedings. ICRA’04. 2004 IEEE International Conference on, vol. 5.IEEE, 2004, pp. 4393–4398.

[9] D. P. Raymer, Aircraft design: a conceptual approach and Rds-student, software for aircraft design, sizing, and performance set(AIAA Education). AIAA (American Institute of Aeronautics & Ast,2006.

[10] J. D.Anderson, Aircraft performance and design. McGraw-Hill, 1999.[11] C. E. Jobe, “Prediction of aerodynamic drag.” DTIC Document, Tech.

Rep., 1984.[12] (2008-11-24) 1th uas team rq-5a hunter2008 unmanned air system

2008. [Online]. Available: http://slideplayer.com/slide/6286116/[13] C. E. Jobe, “Prediction and verification of aerodynamic drag, part i:

prediction,” Progress in Astronautics and Aeronautics, edited by CEEugene, vol. 98, 1985.

[14] L. Meier, P. Tanskanen, F. Fraundorfer, and M. Pollefeys, “Pixhawk:A system for autonomous flight using onboard computer vision,” inRobotics and automation (ICRA), 2011 IEEE international conferenceon. IEEE, 2011, pp. 2992–2997.

[15] E. Zurich, “Qgroundcontrol: Ground control station for small air landwater autonomous unmanned systems,” 2013.

[16] S. Park, J. Deyst, and J. How, “A new nonlinear guidance logicfor trajectory tracking,” in AIAA guidance, navigation, and controlconference and exhibit, 2004, p. 4900.

[17] D. Mellinger and V. Kumar, “Minimum snap trajectory generation andcontrol for quadrotors,” in Proc. of the IEEE Intl. Conf. on Robot. andAutom., Shanghai, China, May 2011, pp. 2520–2525.

[18] (2017-02-11) Px4 firmware. [Online]. Available: https://github.com/PX4/Firmware/tree/master/src/modules

[19] J. Roskam, Airplane flight dynamics and automatic flight controls.DARcorporation, 1995.

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